CN109656149B - Satellite-rocket coupled multi-body system dynamics calculation test method and system - Google Patents
Satellite-rocket coupled multi-body system dynamics calculation test method and system Download PDFInfo
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Abstract
The invention provides a satellite-rocket coupled multi-body system dynamics calculation test method and a system, comprising the following steps: physical mechanical model conversion and topological structure establishment; establishing a transmission matrix and a transmission equation of each element, splicing a total transmission matrix and a total transmission equation of the system by applying a topological structure and a complementary condition, decoupling the system by proving orthogonality analysis, and analyzing the coupling dynamic characteristics of the satellite and the arrow by a modal analysis theory to obtain the inherent characteristics of the system; according to the inherent characteristics of the system, carrying out satellite-rocket coupling dynamic response calculation; and (4) obtaining the stress condition of the satellite-rocket connection boundary through simulation calculation, obtaining the force spectrum input condition of force limit control, and performing a force limit test. According to the method, the satellite-rocket coupling dynamic response is quickly and accurately obtained, the satellite-rocket interface force spectrum is further obtained, the force limit vibration test input conditions are obtained, and the force limit vibration test verification can be quickly and accurately carried out.
Description
Technical Field
The invention relates to the technical field of environmental simulation, in particular to a satellite-rocket coupled multi-body system dynamics calculation test method and system.
Background
The spacecraft dynamics environment mainly occurs in the launching phase, and the importance of the spacecraft dynamics environment is not negligible although the action time is short. Satellites must withstand dynamic loads such as vibration, noise, shock, and acceleration due to various operations such as takeoff, inter-stage (fairing) separation, secondary ignition, shut down, and orbit during launch and powered flight. The dynamic environment can cause structural damage to the spacecraft and components, such as: electronic components are damaged, instrument equipment fails due to short circuit or open circuit of electronic circuits, looseness of connectors and breakage of supports, and a main structure and a secondary structure of the satellite are damaged or broken. The occurrence of these faults may affect the completion of the release mission and even cause the entire mission to fail. In order to ensure that the satellite/spacecraft is not subject to damage when subjected to these harsh environmental conditions, ground mechanics environmental tests must be performed. The satellite vibration test aims to verify the correctness of a satellite structure design scheme and an analysis model for calculation, confirm the rationality of a manufacturing process scheme and provide a basis for determining the design of a sample product.
The force control vibration test is that the output of the vibration table is controlled by monitoring and controlling the interface force between the satellite and the table top clamp and through feedback regulation, a force spectrum control similar to acceleration control is executed. The vibration test eliminates 'over-test' caused by acceleration envelope in the low-frequency stage, and the environment test is closer to the real flight load condition of the active section. Morrow in 1960 proposed mounting a force sensor between the test piece and the fixture to simulate mechanical impedance by controlling the test input force. With the development of force measurement technology, force measuring ring FMD has been developed successfully and applied in NASA and European space Bureau, and applied in vibration tests of complete spacecraft and important loads, such as TOPEX test, Hubble telescope planetary camera, CASSINI detector, etc. According to flight data and theoretical results, NASA in 2013 forms a force control test guide, and a force control test technology becomes an important method for effectively and reasonably solving the problem of over-test.
The spacecraft vibration test condition is optimized, the over-test and under-test problems of the vibration test are avoided as much as possible, the evaluation capability of the vibration test is improved, and the ground test verification requirements of large-scale complex satellites are met. Therefore, the intensive research of the rapid calculation and test method of the dynamics of the satellite-rocket coupled multi-body system becomes a key subject which needs to be faced urgently and solved.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a satellite-rocket coupled multi-body system dynamics calculation test method and system.
The invention provides a satellite-rocket coupled multi-body system dynamics calculation test method, which comprises the following steps:
satellite-rocket coupling multi-body dynamics modeling step: the method comprises the steps of physical mechanical model conversion and topological structure establishment;
analyzing coupling dynamics characteristics of stars and arrows: establishing a transmission matrix and a transmission equation of each element, splicing a total transmission matrix and a total transmission equation of the system by applying a topological structure and a complementary condition, decoupling the system by proving orthogonality analysis, and analyzing the coupling dynamic characteristics of the satellite and the arrow by a modal analysis theory to obtain the inherent characteristics of the system;
satellite-arrow coupling dynamic response calculation step: according to the inherent characteristics of the system, carrying out satellite-rocket coupling dynamic response calculation;
and (3) force limit vibration test: and (4) obtaining the stress condition of the satellite-rocket connection boundary through simulation calculation, obtaining the force spectrum input condition of force limit control, and performing a force limit test.
Preferably, the physical mechanical model conversion includes: the component members, the connecting parts and the connecting forms of the system are equivalent to the connection relationship among the mechanical elements.
Preferably, the mechanical element comprises a body element and a hinge element, the body element comprising: flexible body elements including rigid body elements, lumped mass elements, spring elements, beam element elements, plate element elements, or cylinder elements; the hinge member includes: elastic hinges, sliding hinges, spherical hinges or fixed hinges.
Preferably, the mechanical elements are numbered, a starting point is selected, all the body elements and the hinge elements are numbered sequentially, if a closed-loop system is met, the closed-loop system is converted into an open-loop system through a supplement condition, and a topological structure is finally established through cutting to obtain a full-system tree diagram.
Preferably, the transfer matrix U and the transfer equation z of each mechanical element are derived through model equivalence and the characteristics of each body element and hinge elementO＝UzII is the number of input point, O is the number of output point, the total transmission matrix U of the system is assembled by the supplementary condition and the supplementary equationallSum total transfer equation Uallzall＝0。
Preferably, the augmented feature vector and the augmented operator are constructed, the augmented feature vector is proved to automatically meet orthogonality through establishing an augmented formula meeting symmetry, then boundary conditions are brought in, a feature equation of the system is obtained through a total transfer equation to obtain inherent frequency of the system, a boundary point state vector is obtained through solving the total transfer equation through the inherent frequency, and each element state vector and the system matrix are obtained through the transfer equation.
Preferably, the satellite-rocket coupling dynamics characteristic analysis converts a partial differential equation into an ordinary differential equation, calculates initial conditions and excitation force under a generalized coordinate, calculates the generalized coordinate of the system, and obtains satellite-rocket coupling dynamics response through the generalized coordinate and the augmented feature vector.
Preferably, satellite-rocket connection interface force is obtained through the satellite-rocket coupling dynamic response, and the acceleration response of the key part of the satellite is obtained through a force limit vibration test of the force measuring device.
The invention provides a satellite-rocket coupled multi-body system dynamics calculation test system, which comprises
The satellite-rocket coupling multi-body dynamics modeling module comprises: the method comprises the steps of physical mechanical model conversion and topological structure establishment;
satellite-rocket coupling dynamics characteristic analysis module: establishing a transmission matrix and a transmission equation of each element, splicing a total transmission matrix and a total transmission equation of the system by applying a topological structure and a complementary condition, decoupling the system by proving orthogonality analysis, and analyzing the coupling dynamic characteristics of the satellite and the arrow by a modal analysis theory to obtain the inherent characteristics of the system;
the satellite-rocket coupling dynamic response calculation module: according to the inherent characteristics of the system, carrying out satellite-rocket coupling dynamic response calculation;
force limit vibration test module: and (4) obtaining the stress condition of the satellite-rocket connection boundary through simulation calculation, obtaining the force spectrum input condition of force limit control, and performing a force limit test.
Preferably, the physical mechanical model conversion includes: the component members, the connecting parts and the connecting forms of the system are equivalent to the connection relationship among the mechanical elements.
Compared with the prior art, the invention has the following beneficial effects:
according to the invention, a multi-body system transfer matrix method is combined with a force-limited vibration test, satellite-rocket coupling dynamic response is rapidly and accurately obtained, a satellite-rocket interface force spectrum is further obtained, force-limited vibration test input conditions are obtained, and force-limited vibration test verification can be rapidly and accurately carried out. The rapid algorithm does not need a system overall dynamics equation, is low in matrix order, flexible in modeling and high in programming, can rapidly and accurately analyze the dynamic response of the satellite-rocket coupled multi-body system, provides test conditions for a satellite-rocket force-limited vibration test, more truly simulates the mechanical environment experienced by a satellite, improves test effectiveness, and meets the technical requirements of satellite-rocket dynamic simulation and test.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a satellite-rocket coupled multi-body system dynamics model according to an embodiment of the invention;
FIG. 2 is a star-arrow coupled multi-body system dynamics topology according to an embodiment of the present invention;
FIG. 3 is a schematic structural diagram of a force-limited vibration testing system according to the present invention;
in the figures, 0, 1, 2.. 48 is the element number; z is a radical ofij(i-0, …, 47; j-1, …,48) is a state vector; 301-vibration table, 302-clamp, 303-force sensor, 304-force measuring ring FMD, 305-acceleration sensor, 306-satellite, 307-charge amplifier, 308-signal processing, 309-vibration controller, 310-power amplifier.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
The invention provides a satellite-rocket coupled multi-body system dynamics calculation test method, which comprises the following steps:
satellite-rocket coupling multi-body dynamics modeling step: the method comprises the steps of physical mechanical model conversion and topological structure establishment;
analyzing coupling dynamics characteristics of stars and arrows: establishing a transmission matrix and a transmission equation of each element, splicing a total transmission matrix and a total transmission equation of the system by applying a topological structure and a complementary condition, decoupling the system by proving orthogonality analysis, and analyzing the coupling dynamic characteristics of the satellite and the arrow by a modal analysis theory to obtain the inherent characteristics of the system;
satellite-arrow coupling dynamic response calculation step: according to the inherent characteristics of the system, carrying out satellite-rocket coupling dynamic response calculation;
and (3) force limit vibration test: and (4) obtaining the stress condition of the satellite-rocket connection boundary through simulation calculation, obtaining the force spectrum input condition of force limit control, and performing a force limit test.
The star-arrow coupled multi-body dynamics modeling mainly comprises the steps of establishing a star-arrow coupled dynamics model, breaking the whole into zero, establishing element transmission matrixes and transmission equations, establishing a system total transmission matrix and total transmission equations by an assembly system, verifying orthogonality, analyzing dynamics characteristics, calculating dynamics response and the like. The satellite-rocket coupling dynamic model comprises the steps of converting a physical model into a physical model (shown in figure 1), converting the physical model into a mathematical topological model (shown in figure 2) and the like. The "breaking up to zero" includes establishing a transfer matrix and a transfer equation for each individual element and hinge element, etc. The 'assembly system' comprises the step of 'assembling' the transmission matrix and the transmission equation of each element into the total transmission matrix and the transmission equation of the system. The orthogonality proving and dynamic characteristic analysis comprises that orthogonality analysis is carried out through an augmented feature vector and an augmented operator, and inherent characteristics of the system are obtained through modal analysis. In order to quickly and accurately perform dynamics analysis, for a linear system, the dynamics response of a multi-body system can be obtained by solving an ordinary differential equation; for a nonlinear system, a kinetic equation is linearized, and a kinetic response is obtained through a numerical integration algorithm. And obtaining the satellite-rocket interface force, and obtaining the force control acceleration response through a force limit vibration test (as shown in figure 3). In the force-limited vibration test, a satellite 306 is mounted on a vibration table 301 through a clamp 302 and a force measuring ring FMD304, a force signal arithmetic unit 308 transmits the force signal to a data acquisition unit 309, an acceleration sensor 305 acquires an acceleration signal and transmits the acceleration signal to the data acquisition unit 309 through a charge amplifier, and the vibration table 301 is controlled to vibrate through a power amplifier 310.
Further, the satellite-rocket connection interface force spectrum is obtained through satellite-rocket coupling dynamic response calculation and is used as an input condition of a force limit vibration test, and force signals are controlled through a force ring.
The method does not need a system overall dynamics equation, is low in matrix order, flexible in modeling and high in programming, can quickly and accurately analyze the dynamic response of the satellite-rocket coupled multi-body system, obtains satellite-rocket interface force, provides test conditions for force-limited vibration, and meets the technical requirements of satellite-rocket dynamic simulation and test.
The physical mechanical model conversion comprises the following steps: the component members, the connecting parts and the connecting forms of the system are equivalent to the connection relationship among the mechanical elements. The mechanical element comprises a body element and a hinge element, the body element comprises: flexible body elements including rigid body elements, lumped mass elements, spring elements, beam element elements, plate element elements, or cylinder elements; the hinge member includes: elastic hinges, sliding hinges, spherical hinges or fixed hinges. Numbering mechanical elements such as elements 0, 1, 2.. 48 in fig. 1 and 2, selecting a starting point, numbering all body elements and hinge elements in sequence, if a closed-loop system is met, converting the closed-loop system into an open-loop system through a supplement condition, and finally establishing a topological structure through cutting to obtain a full-system tree diagram.
The topological structure, if the topological structure is a chain system, the total transmission matrix is Uall＝UnUn-1…U1The total transfer equation is zn,0＝Uallz0,1. And the closed-loop system is processed into a bifurcation system and a supplement condition, and finally the tree-shaped topological structure is obtained to carry out transfer matrix and transfer equation derivation.
In fig. 1, 5, 9, 16, 29, 36, and 39 are space rigid bodies, 2, 4, 8, 10, and 15+ i (i is 1, 2, 3, and 4), 17, 18, 19, 20, 25, 26, 27, 28, 30, 32, 33, 34, 40, and 41 are space spring dampers, 3, 14, 31, and 37 are thin plates, 6 and 42 are fixed hinges, 7, 12, 21, 22, 23, 24, 38, 44, and 48 are cylinders, 11, 13, and 45 are space sliding hinges, 35 and 46 are space spherical hinges, 34 and 47 are space beams, and 0 is a ground.
Deriving a transfer matrix U and a transfer equation z of each mechanical element through model equivalence and characteristics of each body element and each hinge elementO＝UzII is the number of input point, O is the number of output point, the total transmission matrix U of the system is assembled by the supplementary condition and the supplementary equationallSum total transfer equation Uallzall＝0。
The method comprises the steps of constructing an augmentation feature vector and an augmentation operator, establishing an augmentation formula meeting symmetry, proving that the augmentation feature vector automatically meets orthogonality, further bringing boundary conditions, obtaining a feature equation of a system through a total transfer equation, obtaining natural frequency of the system, solving the total transfer equation through the natural frequency to obtain a boundary point state vector, and obtaining state vectors and system matrix of each element through the transfer equation.
The satellite-rocket coupling dynamics characteristic analysis converts partial differential equations into ordinary differential equations, calculates initial conditions and exciting force under generalized coordinates, calculates generalized coordinates of a system, obtains satellite-rocket coupling dynamics response through the generalized coordinates and the augmented feature vectors, and uses the satellite-rocket coupling dynamics response as input conditions of a force-limited vibration test.
Satellite-rocket connection interface force is obtained through satellite-rocket coupling dynamic response, force limit vibration test is carried out through a force measuring device, and acceleration response of key parts of the satellite is obtained.
And establishing a state vector through the connection relation between the displacement of the connection point in the system and the internal force, and rewriting the state vector relation of each element into a matrix form to obtain an element transfer matrix. The system total transfer matrix and the transfer equation are used for assembling elements according to the system connection relation. The chain system can obtain a total transmission matrix only by multiplying the transmission matrixes of the elements in sequence; in the closed-loop system, a tree-shaped topological structure and a supplementary equation are required to be simultaneously established to obtain a total transmission equation and a transmission matrix of the system.
On the basis of the satellite-rocket coupled multi-body system dynamics calculation test system, the invention also provides a satellite-rocket coupled multi-body system dynamics calculation test system, which comprises
The satellite-rocket coupling multi-body dynamics modeling module comprises: the method comprises the steps of physical mechanical model conversion and topological structure establishment;
satellite-rocket coupling dynamics characteristic analysis module: establishing a transmission matrix and a transmission equation of each element, splicing a total transmission matrix and a total transmission equation of the system by applying a topological structure and a complementary condition, decoupling the system by proving orthogonality analysis, and analyzing the coupling dynamic characteristics of the satellite and the arrow by a modal analysis theory to obtain the inherent characteristics of the system;
the satellite-rocket coupling dynamic response calculation module: according to the inherent characteristics of the system, carrying out satellite-rocket coupling dynamic response calculation;
force limit vibration test module: and (4) obtaining the stress condition of the satellite-rocket connection boundary through simulation calculation, obtaining the force spectrum input condition of force limit control, and performing a force limit test.
Those skilled in the art will appreciate that, in addition to implementing the system and its various devices, modules, units provided by the present invention as pure computer readable program code, the system and its various devices, modules, units provided by the present invention can be fully implemented by logically programming method steps in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers and the like. Therefore, the system and various devices, modules and units thereof provided by the invention can be regarded as a hardware component, and the devices, modules and units included in the system for realizing various functions can also be regarded as structures in the hardware component; means, modules, units for performing the various functions may also be regarded as structures within both software modules and hardware components for performing the method.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (10)
1. A satellite-rocket coupled multi-body system dynamics calculation test method is characterized by comprising the following steps:
satellite-rocket coupling multi-body dynamics modeling step: the method comprises the steps of physical mechanical model conversion and topological structure establishment;
analyzing coupling dynamics characteristics of stars and arrows: establishing a transmission matrix and a transmission equation of each element, splicing a total transmission matrix and a total transmission equation of the system by applying a topological structure and a complementary condition, decoupling the system by proving orthogonality analysis, and analyzing the coupling dynamic characteristics of the satellite and the arrow by a modal analysis theory to obtain the inherent characteristics of the system;
satellite-arrow coupling dynamic response calculation step: according to the inherent characteristics of the system, carrying out satellite-rocket coupling dynamic response calculation;
and (3) force limit vibration test: and (4) obtaining the stress condition of the satellite-rocket connection boundary through simulation calculation, obtaining the force spectrum input condition of force limit control, and performing a force limit test.
2. The satellite-rocket coupled multi-body system dynamics calculation test method according to claim 1, wherein the physical-mechanical model conversion comprises: the component members, the connecting parts and the connecting forms of the system are equivalent to the connection relationship among the mechanical elements.
3. The satellite-rocket coupled multi-body system dynamics calculation test method according to claim 2, wherein the mechanics elements comprise a body element and a hinge element, the body element comprises: flexible body elements including rigid body elements, lumped mass elements, spring elements, beam element elements, plate element elements, or cylinder elements; the hinge member includes: elastic hinges, sliding hinges, spherical hinges or fixed hinges.
4. The satellite-rocket coupled multi-body system dynamics calculation test method according to claim 3, characterized in that, the elements of the mechanics elements are numbered, the starting point is selected, all the body elements and the hinge elements are numbered sequentially, if a closed-loop system is encountered, the closed-loop system is converted into an open-loop system through complementary conditions, and a topological structure is finally established through cutting, so that a full-system tree diagram is obtained.
5. The satellite-rocket coupled multi-body system dynamics calculation test method according to claim 2, wherein the transfer matrix U and the transfer equation z of each mechanical element are derived through model equivalence and the characteristics of each body element and hinge elementO＝UzII is the number of input point, O is the number of output point, the total transmission matrix U of the system is assembled by the supplementary condition and the supplementary equationallSum total transfer equation Uallzall＝0。
6. The satellite-rocket coupled multi-body system dynamics calculation test method according to claim 1, characterized in that an augmented feature vector and an augmented operator are constructed, the augmented feature vector is proved to automatically meet orthogonality by establishing an augmented formula meeting symmetry, then boundary conditions are brought in, a feature equation of the system is obtained through a total transfer equation, natural frequency of the system is obtained, a total transfer equation is solved through the natural frequency to obtain a boundary point state vector, and state vectors and system matrix types of each element are obtained through the transfer equation.
7. The satellite-rocket coupled multi-body system dynamics calculation test method according to claim 6, wherein the satellite-rocket coupled dynamics characteristic analysis converts partial differential equations into ordinary differential equations, calculates initial conditions and exciting forces under generalized coordinates, calculates the generalized coordinates of a system, and obtains satellite-rocket coupled dynamics responses through the generalized coordinates and the augmented feature vectors.
8. The satellite-rocket coupled multi-body system dynamics calculation test method according to claim 7, wherein satellite-rocket coupled interface force is obtained through the satellite-rocket coupled dynamics response, and the acceleration response of the satellite key part is obtained through a force limit vibration test performed by a force measuring device.
9. A satellite-rocket coupled multi-body system dynamics calculation test system is characterized by comprising
The satellite-rocket coupling multi-body dynamics modeling module comprises: the method comprises the steps of physical mechanical model conversion and topological structure establishment;
satellite-rocket coupling dynamics characteristic analysis module: establishing a transmission matrix and a transmission equation of each element, splicing a total transmission matrix and a total transmission equation of the system by applying a topological structure and a complementary condition, decoupling the system by proving orthogonality analysis, and analyzing the coupling dynamic characteristics of the satellite and the arrow by a modal analysis theory to obtain the inherent characteristics of the system;
the satellite-rocket coupling dynamic response calculation module: according to the inherent characteristics of the system, carrying out satellite-rocket coupling dynamic response calculation;
force limit vibration test module: and (4) obtaining the stress condition of the satellite-rocket connection boundary through simulation calculation, obtaining the force spectrum input condition of force limit control, and performing a force limit test.
10. The satellite-rocket coupled multi-body system dynamics computational test system according to claim 9, wherein the physical-mechanical model transformation comprises: the component members, the connecting parts and the connecting forms of the system are equivalent to the connection relationship among the mechanical elements.
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| CN112326165B (en) * | 2020-09-30 | 2022-12-27 | 北京空间飞行器总体设计部 | Method for acquiring force limit condition of vibration test of satellite and component based on interface force spectrum |
| CN112486196B (en) * | 2020-12-02 | 2022-03-01 | 哈尔滨工业大学 | Aircraft rapid trajectory optimization method meeting strict time and position constraints |
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